WO2014204526A2 - Famille de moteurs à double flux à engrenages - Google Patents

Famille de moteurs à double flux à engrenages Download PDF

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Publication number
WO2014204526A2
WO2014204526A2 PCT/US2014/015500 US2014015500W WO2014204526A2 WO 2014204526 A2 WO2014204526 A2 WO 2014204526A2 US 2014015500 W US2014015500 W US 2014015500W WO 2014204526 A2 WO2014204526 A2 WO 2014204526A2
Authority
WO
WIPO (PCT)
Prior art keywords
fan
gear
turbine
engines
low pressure
Prior art date
Application number
PCT/US2014/015500
Other languages
English (en)
Other versions
WO2014204526A3 (fr
Inventor
Karl HASEL
Jessica TSAY
Original Assignee
United Technologies Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Publication of WO2014204526A2 publication Critical patent/WO2014204526A2/fr
Publication of WO2014204526A3 publication Critical patent/WO2014204526A3/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/40—Transmission of power
    • F05D2260/403—Transmission of power through the shape of the drive components
    • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing

Definitions

  • This application relates to a method of utilizing a single gas turbine engine propulsor core with distinct gear drives and fan diameters.
  • Gas turbine engines are known and, when utilized on aircraft, typically, include a fan delivering air into a bypass duct as propulsion air and also delivering air into a compressor.
  • a method of configuring a plurality of substantially mutually distinct gas turbine engines including the steps of providing each of the engines with respective ones of a plurality of substantially mutually distinct propulsors, each including a fan and a fan drive gear architecture.
  • Each of the engines with respective ones of a plurality of substantially mutually alike gas generators provide each with a compressor section, combustor, and a turbine section configured to drive at a common speed, the compressor section and the fan drive gear architecture.
  • the plurality of mutually distinct propulsors have a mutually distinct diameter fan and/or fan drive architecture with distinct gear reduction ratio.
  • a gas generator has a compressor section with first and second compressors.
  • the turbine section has a first and second turbine.
  • the second turbine drives the first compressor and the fan drive gear architecture.
  • the plurality of mutually distinct gas turbine engines produce substantially the same thrust.
  • the mutually distinct gas turbine engines capable of producing mutually distinct thrust.
  • the plurality of substantially mutually distinct gas turbine engines produce a common overall pressure ratio.
  • a method of configuring a plurality of substantially mutually distinct gas turbine engines includes the steps of providing each of the engines with respective ones of a plurality of substantially mutually distinct gas generators, each with a compressor section, combustor, and a turbine section.
  • Each of the engines is provided with respective ones of a plurality of substantially mutually alike propulsors, each including a fan and a fan drive gear architecture.
  • the turbine section is configured to drive the compressor section and the fan drive gear architecture at a common speed.
  • a method of providing a family of gas turbine engines includes the steps of designing a gas turbine engine, a fan having a first diameter, a first gear reduction having a first gear ratio, a low pressure compressor, a high pressure compressor, a combustor, a low pressure turbine, and a high pressure turbine.
  • the low pressure turbine is designed to drive the low pressure compressor at a common speed and to drive the fan through the gear reduction.
  • a distinct gas turbine engine is provided by selecting a second gear ratio for a second gear reduction, and a second fan diameter for a fan, and placing the second gear reduction having the second gear ratio and the fan rotor having the second fan diameter into the distinct gas turbine engine, such that the low pressure turbine now drives the second gear reduction having the second gear ratio and, in turn, the fan having the second fan diameter.
  • the second fan is less than the first fan diameter and the second gear ratio is higher than the first gear ratio.
  • a family of gas turbine engines has a low pressure compressor to be driven at a common speed with a low pressure turbine and an output shaft also driven by the low pressure turbine.
  • a high pressure compressor is driven by a high pressure turbine, and a combustor.
  • a plurality of gear reductions has different gear ratios to be driven by the output shaft.
  • a plurality of fan rotors has different diameters to be driven by a selected gear reduction.
  • a particular gear reduction is selected in combination with the selection of a particular fan rotor diameter.
  • Figure 1 schematically shows a gas turbine engine.
  • Figure 2A shows a first embodiment of a method according to this invention.
  • Figure 2B shows a second embodiment.
  • Figure 3 shows yet another embodiment.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section
  • turbofan gas turbine engine in the disclosed non- limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.
  • the engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10)
  • the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the engine 20 bypass ratio is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5:1.
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other geared gas turbine engines.
  • the fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet.
  • the flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.
  • “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)] 0'5 .
  • the "Low corrected fan tip speed” as disclosed herein according to one non- limiting embodiment is less than about 1150 ft / second.
  • a gas turbine engine 100 has a core 102, which includes the low pressure compressor 44, the high pressure compressor 52, the combustor 26, the high pressure turbine 54 and the low pressure turbine 46, such as in the Figure 1 engine.
  • Output shaft 104 is a shaft driven by the low pressure turbine 46.
  • the gear reduction 106 is selected in combination with a diameter Di of a fan 108, such that the rotational speed of the fan 108 is as desired given the input speed of the shaft 104.
  • gear ratios equal to or between about 2.4 to 4.2 can be utilized.
  • bypass ratios greater than or equal to 6, and greater than or equal to 8, and greater than or equal to 10, and with pressure ratios of less than or equal to 1.52 may be achieved.
  • Overall pressure ratios of greater than 30: 1, and even greater than or about 50: 1 may be achieved.
  • a gear reduction 112 is selected in combination with the diameter D 2 and a similar or different fan pressure ratio of a fan 114 to achieve a speed for the fan which is as desired for better efficiency and noise control.
  • the diameter D 2 of the fan 114 is smaller than the diameter of the fan 108, providing a smaller bypass ratio.
  • the gear reduction ratio of gear 112 would generally be less than the gear ratio of the gear reduction 106.
  • a disclosed method of configuring a plurality of substantially mutually distinct gas turbine engines 100,110 includes providing each of the engines with respective ones of a plurality of substantially mutually distinct propulsors, each including a fan and a fan drive gear architecture 108/106 and 114/112.
  • Each of the engines 100, 110 is provided with respective ones of a plurality of substantially mutually alike gas generators 102, each with a compressor section, combustor, and a turbine section configured to drive at a common speed the compressor section and the fan drive gear architecture.
  • the plurality of mutually distinct propulsors have a mutually distinct diameter fan and/or fan drive architecture with distinct gear reduction ratio.
  • FIG. 3 shows another aspect, where an engine 200 can be constructed with a single propulsor 202 having a fan and a fan drive gear architecture, which can be associated with a family of distinct gas generators 204a and 204b. The benefits mentioned above would be similar from this arrangement.
  • Figures 2A/2B and Figure 3 can be utilized to provide engines having a common thrust across the family, or having distinct thrust.

Abstract

La présente invention concerne un procédé de configuration d'une pluralité de moteurs à turbine à gaz sensiblement mutuellement distincts comprenant les étapes consistant à doter chacun des moteurs de propulseurs respectifs d'une pluralité de propulseurs sensiblement mutuellement distincts. Chacun comprend un ventilateur et une architecture d'engrenage d'entraînement de ventilateur. Chacun des moteurs est doté de générateurs respectifs d'une pluralité de générateurs de gaz sensiblement mutuellement semblables, chacun étant doté d'une section de compression, d'une chambre de combustion et d'une section turbine configurée pour entraîner à une vitesse commune. La pluralité de propulseurs mutuellement distincts possèdent un ventilateur de diamètre sensiblement distinct et/ou une architecture d'entraînement de ventilateur à rapport de démultiplication distinct.
PCT/US2014/015500 2013-02-17 2014-02-10 Famille de moteurs à double flux à engrenages WO2014204526A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201361765738P 2013-02-17 2013-02-17
US61/765,738 2013-02-17

Publications (2)

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WO2014204526A2 true WO2014204526A2 (fr) 2014-12-24
WO2014204526A3 WO2014204526A3 (fr) 2015-03-05

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3623600A1 (fr) * 2018-09-11 2020-03-18 Pratt & Whitney Canada Corp. Moteur à turbine à gaz et son procédé de création de classes
US11578663B2 (en) 2018-09-11 2023-02-14 Pratt & Whitney Canada Corp. Engine family platform design

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7021042B2 (en) * 2002-12-13 2006-04-04 United Technologies Corporation Geartrain coupling for a turbofan engine
US7832193B2 (en) * 2006-10-27 2010-11-16 General Electric Company Gas turbine engine assembly and methods of assembling same
US8459035B2 (en) * 2007-07-27 2013-06-11 United Technologies Corporation Gas turbine engine with low fan pressure ratio
US8277174B2 (en) * 2007-09-21 2012-10-02 United Technologies Corporation Gas turbine engine compressor arrangement
US9995174B2 (en) * 2010-10-12 2018-06-12 United Technologies Corporation Planetary gear system arrangement with auxiliary oil system

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3623600A1 (fr) * 2018-09-11 2020-03-18 Pratt & Whitney Canada Corp. Moteur à turbine à gaz et son procédé de création de classes
US11578663B2 (en) 2018-09-11 2023-02-14 Pratt & Whitney Canada Corp. Engine family platform design
US11629665B2 (en) 2018-09-11 2023-04-18 Pratt & Whitney Canada Corp. Gas turbine engine and method of creating classes of same

Also Published As

Publication number Publication date
WO2014204526A3 (fr) 2015-03-05

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